Shrinking shock strut system for retractable landing gear

ABSTRACT

A shrinking shock strut system for an aircraft landing gear having a retract actuator that is moveable in length to deploy or retract the landing gear, that includes a shrink strut and a transfer device. The shrink strut may be compressed in length for stowage in the fuselage. The transfer device may be in closed fluid communication with the strut shrink for transferring and receiving hydraulic fluid to and from the strut shrink. When actuated by an aircraft hydraulic system independent of any motion of the retract actuator, the transfer device may drive hydraulic fluid to the strut shrink thereby compressing or shrinking the shrink strut to a partially compressed length.

FIELD OF THE INVENTION

This invention relates in general to aircraft landing gear, and moreparticularly, to a shrinking shock strut system that causes a strut toshrink in length independent of the movement of a landing gear actuatorthat retracts the landing gear into the fuselage of the aircraft.

BACKGROUND

In the design and manufacture of aircraft, it is generally desirable tominimize the space required by aircraft components. One approach tosaving space is to shorten the length of the landing gear before orduring retraction into a stowed position in the aircraft. The shortenedlength of the landing gear may be required by the initial design of anaircraft or may be desired to minimize design changes in futuregenerations of an existing aircraft.

A prior art design for shortening the overall length of a strut duringretraction is described in U.S. Pat. No. 5,908,174 to Churchill et al.The patent discloses a shrinking shock strut system that automaticallyshrinks the length of an aircraft landing gear strut during the landinggear's retraction into the aircraft. The shrinking shock strut systemincludes a hydraulic transfer cylinder that transfers hydraulic fluidinto a strut shrink chamber while simultaneously driving pressurized gasfrom a gas spring chamber of the strut into the hydraulic transfercylinder. The hydraulic transfer cylinder is physically integrated witha landing gear retract actuator such that a linear motion of a retractactuator effects an equal linear motion of the hydraulic transfercylinder. In other words, the hydraulic transfer cylinder ismechanically linked to the landing gear retract actuator such that thehydraulic transfer cylinder cannot operate independently from theretract actuator, and thus the extension and shrinking of the landinggear strut is automatically effected during landing gear retraction anddeployment. On deployment, a hydraulic lock on the transfer fluid in thestrut shrink chamber is removed and the pressurized gas that wastransferred to the transfer cylinder drives the transfer fluid from thestrut shrink chamber back into the transfer cylinder. As the transferfluid exits the strut shrink chamber, the pressurized gas returns to thestrut from the transfer cylinder and causes the strut to return to itsextended length.

Other prior art designs provide a shrink actuator that is independent ofthe retract actuator but require a heavy and cumbersome mechanicallinkage to exert an axial shrinking force on the strut sufficient toovercome the strut's internal pressurized gas bias and cause the strutto shrink. Such designs generally require high hydraulic fluid flowrates that may not be available from the aircraft's high pressure/lowflow rate hydraulic system. Such designs may further require a largerlanding gear envelope in the fuselage of the aircraft to accommodate themechanical linkage and the shrink actuator.

SUMMARY OF THE INVENTION

One aspect of the invention provides a shrinking shock strut system foruse with aircraft landing gear that have a retract actuator, wherein theshrinking shock strut system operates independently from movement of theretract actuator. The shrinking shock strut system receives its powerfrom an aircraft's hydraulic supply, but maintains a separation betweenthe system's closed hydraulic circuit and the aircraft's hydraulicsupply. Using a transfer device to “step-down” the hydraulic pressure,the shrinking shock strut system may provide a higher rate of hydraulicfluid flow to a shrink strut than that available from the aircraft'shydraulic supply. The shrinking shock strut system may be sufficientlycompact and avoid the use of heavy mechanical linkage components usedsolely for shrinking the landing gear.

According to another aspect of the invention, a shrinking shock strutsystem may be provided for an aircraft landing gear having a retractactuator that is moveable in length to deploy or retract the landinggear that includes a shrink strut and a transfer device. The shrinkstrut may be moveable between an extended length and a shrunk length,and may include a strut cylinder and a strut piston mounted coaxially,wherein a strut shrink chamber is formed between the strut cylinder andthe strut piston, and wherein an increase in the length of the strutshrink chamber causes a decrease in the length of the shrink strut. Thesystem may include a transfer device configured to transfer hydraulicfluid to the strut shrink chamber independent of movement of the retractactuator, and may further include a control device configured to directaircraft hydraulic fluid to the transfer device to cause it to transferhydraulic fluid to the strut shrink chamber.

According to another aspect of the invention, the transfer device mayinclude a main cylinder defining a cylindrical chamber and a transferpiston mounted for reciprocal motion within the cylindrical chamber anddividing the cylindrical chamber into a transfer chamber and a supplychamber. The transfer piston may include a transfer pressure area thatis in communication with the transfer chamber and a supply pressure areathat is in communication with the supply chamber, wherein the transferpressure area is larger than the supply pressure area.

According to another aspect of the invention, the shrinking shock strutsystem may use the aircraft's hydraulic supply flow rate in one conduitto drive a higher hydraulic fluid flow rate in a second closed systemconduit.

According to another aspect of the invention, the transfer device canhydraulically lock the flow of hydraulic fluid between the transferdevice and the strut thereby preventing the strut from returning to itsextended length until desired. When commanded, the hydraulic lock on thetransfer device can be released and the gas spring pressure of theshrink strut can be used to drive the fluid from the shrink chamber ofthe strut back into the transfer device.

According to another aspect of the invention, a hydraulic compensatormay be provided to maintain a minimum pressure on the hydraulic fluid inthe shrinking shock strut system and compensate for volume fluctuationsin the hydraulic fluid due to leakage and variations in operatingtemperature. The compensator also may provide an indication in the eventthe volume of hydraulic fluid in the system becomes low.

According to another aspect of the invention, a method is provided forconveying hydraulic fluid into the strut for shrinking the strut.

According to another aspect of the invention, the transfer device mayinclude a transfer cylinder including a transfer piston configured witha small effective pressure area on a hydraulic supply side of thetransfer piston and a large effective pressure area on a transfer sideof the transfer piston. So configured, a small high-pressure flow on thesupply side of the transfer piston can drive a large low-pressure flowfrom the transfer side of the transfer piston through a conduit and intothe shrink strut. The transfer piston may be sealed against the innerwall of the transfer cylinder to separate and isolate the aircraft'shydraulic supply fluid from the shrinking shock strut system's volume oftransfer fluid. A hydraulic supply and valving mechanism may be providedon the aircraft for supplying and directing hydraulic supply fluid tooperate the transfer cylinder.

The foregoing and other features of the invention are hereinafter fullydescribed and particularly pointed out in the claims, the followingdescription, and the annexed drawings setting forth in detail one ormore illustrative embodiments of the invention, such being indicative,however, of but one or a few of the various ways in which the principlesof the invention may be employed.

BRIEF DESCRIPTION OF THE DRAWINGS

FIGS. 1A and 1B are a pair of elevation views of an exemplary aircraftlanding gear in accordance with the invention wherein FIG. 1A shows thelanding gear fully extended and FIG. 1B shows the landing gear in ashrunk state.

FIG. 2 is an isometric view of an exemplary strut and transfer cylinderin accordance with the invention.

FIGS. 3A and 3B are a pair of elevation views in cross-section of anexemplary transfer cylinder in accordance with invention wherein FIG. 3Ashows the transfer cylinder when the landing gear is in an extended orunshrunk state and FIG. 3B shows the transfer cylinder when the landinggear is in a shrunk state.

FIG. 4 is an elevation view in cross-section of an exemplary strut inaccordance with the invention shown when the landing gear is in anextended or unshrunk state.

FIGS. 5A and 5B are a pair of elevation views in cross-section of ashrink chamber region of the strut of FIG. 4 wherein FIG. 5A shows thestrut in an extended or unshrunk state and FIG. 5B shows the strut in ashrunk state.

FIG. 6 is a schematic hydraulic diagram of an exemplary landing gearshrinking shock strut system in accordance with the invention forshrinking a single landing gear.

FIGS. 7A and 7B are a pair of elevation views in cross-section of anexemplary compensator in accordance with the invention wherein FIG. 7Ashows the compensator when the shrinking shock strut system is full offluid and FIG. 7B shows the compensator when the shrinking shock strutsystem is low on fluid.

DETAILED DESCRIPTION

Referring now to the drawings, FIGS. 1A and 1B show an aircraft landinggear 10, depicted in a deployed and extended position in FIG. 1A, andincluding a strut 11 constructed in accordance with the invention.Landing gear 10 is of the type that attaches to the structure of anaircraft and is moveable between a deployed position and a retractedposition. In the retracted position landing gear 10 is housed within thefuselage of the aircraft. One skilled in the art will readily understandthat the invention herein disclosed can be adapted for use inconjunction with body mounted or nose mounted landing gear. In addition,the shrinking shock strut system can be included in laterally retractingas well as forward and aft retracting landing gear configurations.

In FIG. 1B, the landing gear 10 is shown in a deployed but shortened or“shrunk” position. In this position, the strut 11 has been shortened inlength by a distance “d.”

Referring now to FIGS. 1A and 1B in greater detail, landing gear 10includes a hydropneumatic strut 11 including a strut piston 12 having anupper portion that is telescoped inside an open end of a strut cylinder13. A wheel 14 is rotatably mounted to a lower portion of the strutpiston 12. A trunnion shaft 15 is attached normally to an upper endportion of the strut 11. Each end of the trunnion shaft 15 includes ashaft journal that is rotatably mounted in a bushing attached to thestructure of the aircraft. The landing gear 11 pivots about the axis ofthe trunnion shaft 15 to retract following takeoff and pivots in theopposite manner to deploy before landing the aircraft.

A trunnion clevis 16 extends from the outer diameter of the strutcylinder 13 providing an attachment point for a landing gear actuator17. The landing gear actuator 17 is pinned at an actuator piston rod 18extending from a first end to the trunnion clevis 16. A second end ofthe landing gear actuator 17 is pinned to the structure of the aircraft.Hydraulic fluid is provided by the aircraft to the landing gear actuator17 to drive the landing gear 10 to the retracted and deployed positions.

To move the landing gear 10 from the deployed position to the retractedposition, the landing gear actuator 17 compresses, driving the landinggear 10 about the axis of the trunnion shaft 15 to the retractedposition. Before or during landing gear retraction, the strut piston 12may be driven (by the shrinking shock strut system) into the strutcylinder 13 the distance “d” before the landing gear 10 is fullyretracted into the aircraft fuselage. The details concerning the systemfor driving the strut piston 12 into the strut cylinder 13 are providedbelow. The landing gear actuator 17 continues to drive the landing gear10 until it is positioned in the fully retracted position.

To move landing gear 10 from the retracted position to the deployedposition, the landing gear actuator 17 extends, driving the landing gear10 about the axis of the trunnion shaft 15 to the deployed position.Initially, as the landing gear 10 deploys, the strut piston 12 mayremain withdrawn within the strut cylinder 13. Once the landing gear 10clears the aircraft surrounding structure, the strut piston 12 may bedriven by the shrinking shock strut system to extend the strut 11 to itsfully extended length.

Turning to FIG. 2, the strut 11 of FIGS. 1A and 1B is shownisometrically with a transfer cylinder 19 secured to the strut cylinder13. The transfer cylinder 19 is in fluid communication with the strut 11via a transfer port 20 on the transfer cylinder, through a first conduit21, and into a transfer inlet port 22 on the strut 11. One advantage ofthe shrinking shock strut system in accordance with the invention isthat the transfer cylinder 19 may be located anywhere on the landinggear or aircraft so long as it can maintain fluid communication with thestrut 11. A second conduit (not shown) provides fluid communicationbetween the transfer cylinder 19 and the aircraft's hydraulic supply(not shown). Also shown in FIG. 2 is a compensator 23.

Referring now to FIGS. 3A and 3B, the transfer cylinder 19 is shown incross section in two states of operation. FIG. 3A shows the transfercylinder 19 in an extended position that corresponds to the strut 11 inan “un-shrunk” state. FIG. 3B shows the transfer cylinder 19 in acompressed position that corresponds to the strut 11 in a “shrunk andlocked” state. Referring to FIG. 3A, the transfer cylinder 19 includes amain cylinder 30 forming a cylindrical chamber therein. A transferpiston 31 is reciprocally disposed within the cylindrical chamber andincludes two effective pressure areas, one on each side of the piston.The transfer piston 31 divides the cylindrical chamber into twochambers, namely, a transfer chamber 32 and a supply chamber 33. Thetransfer chamber 32 may be opened to the atmosphere through a bleedorifice 34, though in operation the transfer chamber 32 is a part of aclosed hydraulic circuit and is filled with hydraulic fluid. The supplychamber 33 is a part of a second hydraulic circuit and is filled withhydraulic fluid provided by the aircraft's hydraulic supply. When thehydraulic supply fluid is provided by the aircraft under pressure to thesupply chamber 33, the applied pressure acts against the effectivepressure area of the transfer piston 31, urging the transfer piston 31to expand the supply chamber 33 and contract the transfer chamber 32.This motion drives the hydraulic fluid out of the transfer chamber 32and into the closed hydraulic circuit. Appropriate seals 35 are providedabout the transfer piston 31 to maintain a separation of the hydraulicfluids in the transfer chamber 32 and the supply chamber 33, as can bereadily appreciated by one skilled in the art.

The transfer chamber 32 is in fluid communication with the strut 11through the closed hydraulic circuit via the transfer port 20 and aconduit 21 (see FIG. 2). As noted above, when the supply chamber 33receives hydraulic fluid from the aircraft hydraulic supply, thetransfer piston 31 moves to contract the transfer chamber 32 therebydisplacing a specific volume of transfer fluid from the transfer chamber32, through the transfer port 20 and into the strut 11 by way of theconduit 21. FIG. 3B shows the transfer cylinder 19 after the transferpiston 31 has displaced the entire volume of transfer fluid from thetransfer chamber 32.

When the transfer piston 31 is positioned within the main cylinder 30such that it has fully displaced the volume of transfer fluid, a lockingmechanism 36 engages to close and lock the transfer piston 31, providinga hydraulic block on the transfer port 20 such that the volume oftransfer fluid cannot reenter the transfer chamber 32. The transferpiston 31 is configured to cooperate with the locking mechanism 36 toeffect the engagement. Those skilled in the art of actuation lockingmechanisms would be able to introduce such a lock and the exemplarysystem does not depend on any particular variety.

Also shown in FIGS. 3A and 3B is an unlock port 37 located at an upperportion of the main cylinder 30 and adjacent the transfer port 20. Theunlock port 37 is in fluid communication with the aircraft's hydraulicpower supply via a conduit (not shown). To unlock and open the transferport 20, a valve (not shown) directs a pressure (P_(unlock)) from theaircraft's hydraulic supply to the unlock port 37 to displace anunlocking pin 38 and release the locking mechanism 36.

The supply chamber 33 is in fluid communication with the aircraft'shydraulic supply via a shrink port 39, and a conduit (not shown). Theshrink port 39 is located on a portion of the main cylinder 30 oppositethe transfer piston 31 from the transfer port 20 such that it is influid communication with the supply chamber 33 and not in fluidcommunication with the transfer chamber 32. The transfer piston 31 andseals 35 separate the transfer chamber 32 from the supply chamber 33such that they are not in fluid communication. This separation preventsthe volume of transfer fluid from entering and affecting the aircraft'shydraulic power supply and prevents the hydraulic power supply fromaffecting the closed hydraulic circuit of the shrinking shock strutsystem.

In operation, the transfer cylinder 19 responds to a relatively highpressure and low flow rate hydraulic fluid input from the aircraft'shydraulic system to transfer or provide a lower pressure and higher flowrate hydraulic fluid output to the strut 11 to reduce the length of thestrut 11. The aircraft's hydraulic supply applies a pressure(P_(supply)) at the shrink port 39 and hydraulic fluid enters into thesupply chamber 33. When the effective force applied to the piston 31from the supply chamber 33 (P_(supply)) is greater than an effectiveforce applied to the piston 31 from the transfer chamber 32(P_(shrink)), the transfer piston 31 moves to increase the size of thesupply chamber 33 and contract the size of the transfer chamber 32,thereby displacing a volume of transfer fluid through the transfer port20 and into the strut 11 to shrink it for retraction into the fuselage.

When the landing gear is again deployed, the aircraft's hydraulic supplymay reduce the pressure applied at the shrink port 39 (now P_(return))and applies an unlock pressure at the unlock port 37 (P_(unlock))causing the locking mechanism 36 to release, thereby removing thehydraulic block on flow through transfer port 20. If the reducedpressure applied at the shrink port 39 (P_(return)) results in aneffective force on the piston 31 that is below the effective forceapplied by a pressure in the closed hydraulic circuit (P_(gas)), thenthe volume of transfer fluid will return to the transfer chamber 32 viathe transfer port 20 and return the transfer piston 31 to its originalposition. The pressures (P_(return)) and (P_(gas)) are discussed belowwith reference to FIG. 6.

Turning now to FIG. 4, a cross-section of the strut 11 is shown in afully extended position as is the case after takeoff but prior toretraction. As can be seen, the transfer cylinder 19 (shown secured tothe strut 11) is in an extended state indicating that the volume oftransfer fluid resides in the transfer chamber 32. The strut piston 12is reciprocally received within the strut cylinder 13 in a concentricrelationship with and between the strut cylinder 13 and an orificesupport tube 40. A lower strut bearing 41 is received within the openend of the strut cylinder 13 against which the strut piston 12 slides.An upper strut bearing 42 is attached to the end portion of the strutpiston 12 within the strut cylinder 13 sliding against its interior. Agas spring chamber 43 is formed within the interior section of thecylinder 13, piston 12, and capped by the lower strut bearing 41. Thegas spring chamber 43 is pressurized with gas at a pressure (P_(gas)),which varies with the temperature and degree of insertion of the strutpiston 12 into the strut cylinder 13. A shrink piston 44 and a set ofspacers 45 are disposed between the interior bore of the strut cylinder13, the strut piston 12, the upper strut bearing 42, and the lower strutbearing 41, allowing the strut piston 12 to extend a predetermineddistance. At full strut extension, the upper strut bearing 42, shrinkpiston 44, a set of spacers 45, and lower strut bearing 41 are forcedinto contact due to the bias caused by the gas pressure (P_(gas)) withinthe gas spring chamber 43.

A strut shrink chamber 46 is formed between the shrink piston 44, thestrut cylinder 13, the strut piston 12, and the lower bearing 41. Thestrut shrink chamber 46 is in fluid communication with the transferinlet port 22 and is filled with hydraulic fluid. In FIG. 4 the strut 11is shown unshrunk. In this state, the shrink chamber 46 has a minimumvolume. The shrink chamber 46 is in fluid communication with thetransfer chamber 32 of the transfer cylinder 19 via a conduit 21 (shownschematically) and the volume of transfer fluid may be transferred fromthe transfer chamber 32 to the shrink chamber 46 when the transferchamber force on the shrink piston 44 exceeds the force applied to theopposite side of the shrink piston 44 (resulting from P_(gas) and otherexternally applied forces).

The gas spring pressure of the strut 11 varies with the compression ofthe piston 12 into the cylinder 13, as in the case when the aircraftlands or is taxiing across the airfield. In such a case, the pressurizedgas in the gas spring chamber 43 is further compressed as the strut 11absorbs the energy of the loads exerted upon it. The shrink piston 44may remain in contact with the lower strut bearing 41 and a reboundchamber 47 is established between the upper strut bearing 42 and theshrink piston 44. As the strut piston 12 again extends, the fluid in therebound chamber 47 slows the piston 12 extension rate with increasedpressure which is exerted against the shrink piston 44 keeping it incontact with the lower bearing 41.

Turning to FIGS. 5A and 5B, detailed cross-sectional views of the strut11 are shown in two states: unshrunk (FIG. 5A) and shrunk (FIG. 5B). InFIG. 5A, the shrink piston 44 is shown disposed between the lowerbearing 41 and the rebound chamber 47. In response to the movement ofthe transfer piston 31, the volume of hydraulic fluid in the transferchamber 32 flows through the transfer inlet port 22 into the strutshrink chamber 46. The fluid pressure in the strut shrink chamber 46 issufficient to force the shrink piston 44, spacers 45, upper strutbearing 42, and strut piston 12 to withdraw into the strut cylinder 13 adistance “d.”

FIG. 5B illustrates a fully shrunk state wherein the shrink chamber 46has expanded to accommodate the volume of transfer fluid transferredfrom the transfer chamber 32. As mentioned previously, once the transferpiston 31 has displaced the volume of fluid from the transfer chamber32, the locking mechanism 36 is engaged thus locking transfer piston 31which blocks flow through transfer port 20 such that the volume oftransfer fluid is hydraulically “locked” in the shrink chamber 46.

Turning now to FIG. 6, a control device such as an exemplary aircraftsequence valve 60 commands the aircraft's hydraulic supply to (1) shrinkthe strut 11 and independently (2) retract the landing gear 10. Thesequence valve 60 which directs hydraulic supply pressure and returnpressure to the transfer cylinder 19 is in fluid communication with thesupply chamber 33 of the transfer cylinder 19 through a conduit 61 andsupply port 39. The sequence valve 60 is also in fluid communicationwith the locking mechanism 36 through a conduit 62 and unlock port 37.As noted above, the conduit 21 provides fluid communication between thestrut shrink chamber 46 and the transfer chamber 32. An appropriatehydraulic supply system (not shown) including, for example, a hydraulicpump and fluid reservoir, is necessary to supply the sequence valve 60.While the example of sequence valve 60 is shown and described, controldevices are well known in the art and the invention is not dependent onthe use of any particular control device.

In accordance with the invention, the shrinking shock strut system iscontrolled independently of the landing gear actuator 17. It is noted,however, that the sequence valve 60 may operate the shrinking shockstrut system described herein before or during landing gear retractionand may return the strut to its original length during or after landinggear deployment.

Applying a numerical example to the schematic in FIG. 6, an aircrafthydraulic supply may be provided at a relatively high pressure (e.g.,4000 psi) but with a relatively low flow rate (e.g., 3.0 GPM) available.Such a low flow rate may not provide the volume of flow required to fillthe shrink chamber 46 of the strut 11 in a given period of time betweenthe pilot's input and the landing gear's retraction into the fuselage ofthe aircraft. When, however, the high supply pressure is provided to thetransfer cylinder 19, it acts on a small effective pressure area of thetransfer piston 31. The opposing side of the transfer piston 31 has amuch larger effective area and—when the piston 31 is in motion—it drivesa much larger transfer fluid flow rate (e.g., 6.5 GPM per landinggear—13 GPM for left and right landing gear) from the transfer chamber32 at a lower pressure (e.g., less than 2000 psi). For returning thestrut 11 to its unshrunk or extended length, the aircraft sequence valve60 ensures that the supply shrink pressure (P_(supply)) is removed orreduced (P_(return)), and further applies an unlocking pressure(P_(unlock)) via conduit 62 to the unlock port 37. This combination ofinputs operates to release the locking mechanism 36 and unblock flowthrough transfer port 20 to allow the gas spring pressure (P_(gas)) ofthe strut to drive the shrink piston 44 to contract the shrink chamber46 and return the volume of transfer fluid to the transfer chamber 32.

The shrinking shock strut system in accordance with the invention relieson transferring a specific volume of transfer fluid from the transferchamber 32 to the shrink chamber 46. Compressibility and in-servicetemperature variations may affect the effective fluid volume in thesystem. To compensate for such fluctuations, the compensator 23 is addedto the system's hydraulic schematic shown in FIG. 6. The compensator 23provides an additional volume of fluid to the system to compensate forthe factors listed above and further maintains a nominal pressure on thefluid in the conduits of the system and provides an indication when thefluid level in the system becomes low and maintenance is required.

Turning to FIGS. 7A and 7B, an exemplary compensator 23 is shown incross-section in two states of operation. FIG. 7A shows the compensator23 when it is full of hydraulic fluid. FIG. 7B shows the compensator 23when the shrinking shock strut system is low on hydraulic fluid. Thecompensator 23 includes a main cylinder 70 that forms a cylindricalcavity having two ends. At a first end of the cavity is a fluid chamber71 that is in fluid communication with the shrinking shock strutsystem's hydraulic circuit via a port 72. Between the fluid chamber 71and a second end of the cylindrical cavity is a piston 73 that maintainsa seal against an interior wall of the cylindrical cavity. On a side ofthe piston 73 opposite the fluid chamber 71 is a spring 74 exerting apreloaded bias force against piston 73 for maintaining a nominalpressure on the fluid in the fluid chamber 71. In the example shown thespring is a mechanical coil spring. Any spring, however, that provides abias force against the piston 73 and thus a pressure in the fluid in thefluid chamber 71 may be used. Examples include a compressed gas springor a different mechanical spring configuration.

In the example shown in FIGS. 7A and 7B, the compensator 23 alsoincludes a low fluid visual indicator 75 by way of adding an extension76 to the piston 73 that extends through an orifice 77 at the first endof the cylindrical cavity when the spring 74 pushes the piston 73through the fluid chamber 71. The presence of the indicator 75 throughthe orifice 77 indicates that the fluid chamber 71 is marginally filledor empty. This visual indication provides the aircraft maintenance crewa simple check as to whether the shrinking shock strut system has lostfluid or otherwise requires maintenance. A sensor could also be added toprovide an electronic indication to a maintenance computer along with(or instead of) the visual indication.

It is further noted that the exemplary compensator 23 may be packagedand located anywhere in the aircraft, the transfer cylinder 19, or thestrut 11.

Although the invention has been shown and described with respect to acertain preferred embodiment or embodiments, it is obvious thatequivalent alterations and modifications will occur to others skilled inthe art upon the reading and understanding of this specification and theannexed drawings. In particular regard to the various functionsperformed by the above described elements (components, assemblies,devices, compositions, etc.), the terms (including a reference to a“means”) used to describe such elements are intended to correspond,unless otherwise indicated, to any element which performs the specifiedfunction of the described element (i.e., that is functionallyequivalent), even though not structurally equivalent to the disclosedstructure which performs the function in the herein illustratedexemplary embodiment or embodiments of the invention. In addition, whilea particular feature of the invention may have been described above withrespect to only one or more of several illustrated embodiments, suchfeature may be combined with one or more other features of the otherembodiments, as may be desired and advantageous for any given orparticular application.

1. A shrinking shock strut system for an aircraft landing gear having aretract actuator that is moveable in length to deploy or retract thelanding gear, comprising: a shrink strut that is moveable between anextended length and a shrunk length, the shrink strut including a strutcylinder and a strut piston mounted coaxially, wherein a variable lengthstrut shrink chamber is formed between the strut cylinder and the strutpiston, and wherein an increase in the length of the strut shrinkchamber causes a decrease in the length of the shrink strut; a transferdevice configured to transfer hydraulic fluid to the strut shrinkchamber independent of movement of the retract actuator; a controldevice configured to cause the transfer device to transfer hydraulicfluid to the strut shrink chamber; and a pressure compensator andwherein the strut shrink chamber, the transfer device, and the pressurecompensator are in communication with a closed hydraulic system, whereinthe pressure compensator is configured to maintain a minimum pressure inthe closed hydraulic system.
 2. The shrinking shock strut systemaccording to claim 1, wherein the transfer device comprises: a maincylinder defining a cylindrical chamber therein; and a transfer pistonmounted for reciprocal motion within the cylindrical chamber anddividing the cylindrical chamber into a transfer chamber and a supplychamber; wherein the transfer piston comprises a transfer pressure areathat is in communication with the transfer chamber and a supply pressurearea that is in communication with the supply chamber; and wherein thetransfer pressure area is larger than the supply pressure area.
 3. Theshrinking shock strut system according to claim 1, wherein the pressurecompensator is further configured to provide additional hydraulic fluidto the closed hydraulic system to account for variations in fluidvolume.
 4. The shrinking shock strut system according to claim 1,wherein the pressure compensator comprises: a chamber configured tostore hydraulic fluid under pressure, a piston configured to applypressure to hydraulic fluid stored in the chamber, a spring configuredto exert a biasing force on the piston, and an opening providingcommunication between the chamber and the closed hydraulic systemconfigured for providing and receiving hydraulic fluid from the closedhydraulic system.
 5. The shrinking shock strut system according to claim1, wherein the pressure compensator comprises a low fluid indicator. 6.The shrinking shock strut system according to claim 1, wherein a controldevice comprises a valve for directing aircraft hydraulic supply fluidto the transfer device.
 7. The shrinking shock strut system according toclaim 1, wherein the transfer device comprises a piston having a supplyside and a transfer side, and wherein an effective pressure area on thesupply side is smaller than an effective pressure area on the transferside.
 8. A shrinking shock strut system for an aircraft landing gearhaving a retract actuator that is moveable in length to deploy orretract the landing gear, comprising: a shrink strut that is moveablebetween an extended length and a shrunk length, the shrink strutincluding a strut cylinder and a strut piston mounted coaxially, whereina variable length strut shrink chamber is formed between the strutcylinder and the strut piston, and wherein an increase in the length ofthe strut shrink chamber causes a decrease in the length of the shrinkstrut; a transfer device configured to transfer hydraulic fluid to thestrut shrink chamber independent of movement of the retract actuator,wherein the transfer device includes a main cylinder defining acylindrical chamber therein and a transfer piston mounted for reciprocalmotion within the cylindrical chamber and dividing the cylindricalchamber into a transfer chamber and a supply chamber, wherein thetransfer piston includes a transfer pressure area that is incommunication with the transfer chamber and a supply pressure area thatis in communication with the supply chamber, and wherein the transferpressure area is larger than the supply pressure area; a control deviceconfigured to cause the transfer device to transfer hydraulic fluid tothe strut shrink chamber; a first conduit configured to establish closedfluid communication between the transfer chamber and the strut shrinkchamber; and a second conduit configured to establish fluidcommunication between the supply chamber and an aircraft hydraulicsystem.
 9. The shrinking shock strut system according to claim 8,wherein the transfer device further comprises a locking mechanism. 10.The shrinking shock strut system according to claim 9, wherein thelocking mechanism is configured to hydraulically lock the hydraulicfluid in the first conduit.
 11. The shrinking shock strut systemaccording to claim 10, further comprising a third conduit configured toestablish fluid communication between the locking mechanism and anaircraft hydraulic system.
 12. The shrinking shock strut systemaccording to claim 8, further configured to provide a higher fluid flowrate to the strut shrink chamber through the first conduit than thefluid flow rate it receives from the aircraft supply through the secondconduit.